The present invention relates generally to gas turbine engines, and, specifically, to clearance control between turbine rotor blade tips and a stator shroud spaced radially thereabove.
A gas turbine engine includes in serial flow communication one or more compressors followed in turn by a combustor and high and low pressure turbines disposed axisymmetrically about a longitudinal axial centerline within an annular outer casing. During operation, the compressors are driven by the turbine and compress air which is mixed with the fuel and ignited in the combustor for generating hot combustion gases. The combustion gases flow downstream through the high and low pressure turbines which extract energy therefrom for driving the compressors and producing output power either as shaft power or thrust for powering an aircraft in flight, for example.
Each of the turbines includes one or more stages of rotor blades extending radially outwardly from respective rotor disks, with the blade tips being disposed closely adjacent to a turbine shroud supported from the casing. The tip clearance defined between the shroud and blade tips should be made as small as possible since the combustion gases flowing therethrough bypass the turbine blades and therefore provide no useful work. In practice, however, the tip clearance is typically sized larger than desirable since the rotor blades and turbine shroud expand and contract at different rates during the various operating modes of the engine.
The turbine shroud has substantially less mass than that of the rotor blades and disk and therefore responds at a greater rate of expansion and contraction due to temperature differences experienced during operation. Since the turbines are bathed in hot combustion gases during operation, they are typically cooled using compressor bleed air suitably channeled thereto. In an aircraft gas turbine engine for example, acceleration burst of the engine during takeoff provides compressor bleed air which is actually hotter than the metal temperature of the turbine shroud. Accordingly, the turbine shroud grows radially outwardly at a faster rate than that of the turbine blades which increases the tip clearance and in turn decreases engine efficiency. During a deceleration chop of the engine, the opposite occurs with the turbine shroud receiving compressor bleed air which is cooler than its metal temperature causing the turbine shroud to contract relatively quickly as compared to the turbine blades, which reduces the tip clearance.
Accordingly, the tip clearance is typically sized to ensure a minimum tip clearance during deceleration, for example, for preventing or reducing the likelihood of undesirable rubbing of the blade tips against the turbine shrouds.
The turbine shroud therefore directly affects overall efficiency or performance of the gas turbine engine due to the size of the tip clearance. The turbine shroud additionally affects performance of the engine since any compressor bleed air used for cooling the turbine shroud is therefore not used during the combustion process or the work expansion process by the turbine blades and is unavailable for producing useful work. Accordingly, it is desirable to reduce the amount of bleed air used in cooling the turbine shroud for maximizing the overall efficiency of the engine.
In order to better control turbine blade tip clearances, active clearance control systems are known in the art and are relatively complex for varying during operation the amount of compressor bleed air channeled to the turbine shroud. In this way the bleed air may be provided as required for minimizing the tip clearances, and the amount of bleed air may therefore be reduced. However, in order to minimize the complexity and cost of providing clearance control, typical turbine shrouds are unregulated in cooling the various components thereof.
Furthermore, in order to control the blade tip clearance, flow of the compressor bleed air through the turbine shroud must also be controlled. Uncontrolled leakage of the bleed air through the various Joints in the turbine shroud assembly directly affects heat transfer and therefore thermal performance of the shroud. And, uncontrolled leakage of the bleed air from the shroud cavity disposed directly above each of the shroud panels has an undesirable effect on backflow margin. Backflow margin is a conventional parameter which indicates the pressure gradient across the shroud panels with a higher pressure being desired above the panels relative to the pressure of the combustion gases which flow along the inner surfaces thereof. Unless the backflow margin is maintained at a suitable level, combustion gases could be undesirably ingested backwardly through the various cooling holes provided for discharging the bleed air through the panels. This could considerably shorten the useful life of the shroud panels during operation.